Solid-rocket propellants

ABSTRACT

Solid-fuel rocket propellants comprising an oxidizer, an oxophilic metal-halophilic metal formulation, and a binder are described herein. Further described are processes for preparing such propellants and methods of reducing hydrogen chloride production via the combustion of such propellants. Non-limiting examples of such formulations include aluminum-lithium alloys.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of PCT international applicationnumber PCT/US2016/021370, filed Mar. 8, 2016, which claimed benefit ofU.S. provisional application No. 62/130,108, filed Mar. 9, 2015 and U.S.provisional application No. 62/130,122, filed Mar. 9, 2015, the contentsof which are hereby incorporated by reference in their entirety intothis disclosure.

STATEMENT OF GOVERNMENT INTEREST

This invention was made with government support under FA9550-13-1-0004awarded by the United States Air Force, and 32 CFR 168a awarded by theDepartment of Defense. The government has certain rights in theinvention.

BACKGROUND

The most common oxidizer used in solid rocket propellants is ammoniumperchlorate (“AP”), NH₄ClO₄. This oxidizer is used because of its highoxygen content and large gas volume that is generated during combustion.However, the chlorine in the oxidant generally favors formation ofhydrogen chloride as its primary product species. In common propellantformulations using AP, as much as 98% of the available chloride ion maybe converted to hydrogen chloride. Hydrogen chloride pollution hasserious negative consequences both to the environment and the ozonelayer and is a major contributor to launch site equipment corrosion.Hydrogen chloride also forms nucleation sites for aerosolized waterproducts which contributes to secondary smoke formation from exhaustplumes thereby making rockets easier to detect, which is a disadvantagefor some applications.

There are three methods that have been investigated generally forhydrogen chloride reduction in composite rocket propellants, namelyusing reduced chlorine propellants, neutralizing hydrogen chloride, andthe use of scavengers.

Using reduced or non-chlorine containing oxidizers reduces or eliminateshydrogen chloride formation, though generally at an unacceptable loss inperformance or increase in detonation sensitivity. Ammonium nitrate(“AN”) has been widely investigated as an AP replacement, but itsperformance has been shown to be much lower than AP and with lessballistic tailorability. Others have investigated incorporating nitrateesters and/or nitroamines in place of all or part of the AP oxidizer.While these energetic formulations have been widely used for their highperformance, they are often more sensitive and are considered Class 1.1propellants which make their handling substantially more dangerous.

Acid-base chemistry has been investigated as a way to neutralizehydrogen chloride formed from solid rocket propellants by usingmagnesium as a fuel in propellants. When burned with AP, magnesium oxideand hydrogen gas are formed. Hydrogen gas will subsequently after burnwith ambient oxygen outside of the rocket motor to form water which inturn reacts with the magnesium oxide to form magnesium hydroxide. Themagnesium hydroxide may then react with hydrogen chloride to formmagnesium chloride and water. While this propellant has adequatetheoretical performance and neutralizing effects, it also has a lowdensity thus requiring a heavier loading and a serious disadvantage inthat it requires sufficient atmospheric oxygen to combust hydrogen. Dueto the low levels of oxygen at higher altitudes, this approach haslimited applicability.

Another possible method for reducing hydrogen chloride formation wouldbe to use strongly halophilic materials, such as alkali metals, to“scavenge” chlorine ions during combustion to form alkali-metalchlorides. Because of the high reactivity of these materials andpotential health effects, alkali metals are generally not used as neatelemental metals in formulations. Therefore, alkali metal nitrates, suchas LiNO₃ and NaNO₃ have been used to stably introduce the halophilicmaterial, replacing a stoichiometric amount of AP. While theseformulations do reduce the hydrogen chloride formation duringcombustion, they do so at an unacceptable loss to performance. It hasbeen shown that the addition of nitroamines to the formulation canoffset this performance deficit to some degree, but their additionresult is a Class 1.1 propellant, which is extremely dangerous tohandle.

While some general research into dual oxophilic-halophilic fuelcombinations has been investigated as a means to improve specificimpulse (I_(SP)) when combined with a chlorine-containing oxidant, suchas disclosed in U.S. Pat. No. 3,133,841, such use as hydrogen chloridescavengers has previously not been investigated and the range of alloyspresented lack suitable performance characteristics. There is,therefore, an unmet need for a high performance solid-fuel rocketpropellant capable of significantly reducing hydrogen chlorideformation.

SUMMARY OF INVENTION

In one aspect of the invention, a solid-rocket propellant comprising anAl—Li alloy with a weight ratio of Li to Al between about 14% and 34%,an oxidizer, and a binder is provided.

In another aspect of the invention, a process for reducing hydrogenchloride formation in solid-rocket combustion comprising the steps ofcombining a halophilic metal with an oxophilic metal to form an alloy,combining the alloy with an oxidizer and a binder to form a propellant,and combusting the propellant is provided.

In a further aspect of the invention, a method for producing asolid-form propellant is provided comprising formulating a halophilicmetal with an oxophilic metal to form a plurality of formulatedoxophilic-halophilic metal particles, and combining the formulatedoxophilic-halophilic particles with a chlorine-containing oxidizer and abinding agent to form a solid-form propellant.

In a still further aspect of the invention, a solid-rocket propellant isprovided comprising an oxophilic metal-halophilic metal formulation, abinder, and an oxidant is provided.

In yet another aspect of the invention, a chloride containingsolid-rocket propellant is provided where substantially no hydrogenchloride is produced upon combustion.

In still a further aspect of the invention, a method for reducing largemolten droplets (“LMD”) in solid-rocket fuel propellants is provided.

BRIEF DESCRIPTION OF DRAWINGS

FIG. 1A shows the thermochemical equilibrium results for neataluminum/AP/HTPB propellant formulations. Specific impulse contour linesabove 250 seconds represent a step of 2.5 seconds, contour lines below250 seconds represent a step of 5 seconds, and transverse contour linesrepresent propellant solids loading.

FIG. 1B shows the thermochemical equilibrium results for neatlithium/AP/HTPB propellant formulations. Specific impulse contour linesabove 250 seconds represent a step of 2.5 seconds, contour lines below250 seconds represent a step of 5 seconds, and transverse contour linesrepresent propellant solids loading.

FIG. 1C shows the thermochemical equilibrium results for 80/20aluminum-lithium/AP/HTPB propellant formulations. Specific impulsecontour lines above 250 seconds represent a step of 2.5 seconds, contourlines below 250 seconds represent a step of 5 seconds, and thetransverse contour lines represent propellant solids loading.

FIG. 2 shows the theoretical performance of the Al—Li system as afunction of lithium content.

FIG. 3 shows the conversion of Cl→hydrogen chloride, Cl→LiCl, andLi→LiCl as a function of Li content in the fuel additive.

FIG. 4 is a particle size distribution curve of neat Al and Al—Li.

FIG. 5 are DSC/TGA and ion current curves for neat AP.

FIG. 6 are DSC/TGA and ion current curves for a 30% Aluminum/70% APmixture.

FIG. 7 are DSC/TGA and ion current curves for a 30% Aluminum-Lithiumalloy/70% AP mixture.

FIG. 8 is an HCl deter apparatus.

FIG. 9 are FUR spectra of the decomposition of neat AP and a 30%Al—Li/AP mix.

DETAILED DESCRIPTION

A solid-rocket propellant is provided which comprises a binder, anoxidizer and an oxophilic metal-halophilic metal formulation. Theformulation is selected so as to have the capability to reduce hydrogenchloride formation when a chlorine-containing oxidizer, such as AP isused as an oxidizer, while at the same time maintaining acceptableperformance.

Thermochemical calculations as set forth in Example 1 and Example 2 showthat binary oxophilic-halophilic fuels result in good performance whenreacted with a chorine-containing oxidant. The halophilic fuel can becombined through alloying, mechanical activation, or encapsulation witha suitable oxophilic-metal fuel. The resulting formulatedoxophilic-halophilic metal composition can then be combined with anoxidant and a binder to form a propellant. When combusted, thepropellant results in reduced hydrogen chloride formation and improvedperformance.

By using this formulation approach, a stoichiometricoxophilic:halophilic (metal) ratio can be tailored to minimize hydrogenchloride formation, minimize non-alkali chloride species formation, andmaximize performance. Both modeling and performance testing ofpropellants to illustrate certain embodiments of the invention wereperformed. For example, various aluminum-lithium alloy formulationsshowed both a predicted and experimental reduction of over 95% in theformation of hydrogen chloride gas. Further, modeling indicated apredicted about a greater than 2% increase in motor performance inspecific impulse (I_(sp)), an important performance metric.

These modeling investigations involve theoretical performancecalculations using various Al—Li ratios, as well as neat aluminum, neatlithium, and a baseline scavenged propellant replacing a portion of APwith NaNO₃. This work was completed to determine a range of lithiumcontents for which the Al—Li system could act as a high performancehalide scavenger (“HPHS”), as set forth in Example 2. Examples of theseternary system simulations are presented in FIG. 1. The mixture ratiosthat yielded the highest specific impulse for each fuel additive aretabulated in Table 1 and their corresponding performance values aretabulated in Table 2. The mixture ratio with the highest HPHS value forthe 80/20 Al—Li system is denoted by an asterisk (*) in Tables 1-2, andis the mixture ratio that was used for all experimental results. Thesolids loading differs at the highest HPHS value than the correspondingloading for maximum I_(SP).

The performance values in Table 2 include specific impulse, impulsedensity, enthalpy change across the nozzle, chamber temperature, exhaustmolecular weight, and the conversion of chlorine and alkali metal to HCland MCl where “M” denotes alkali metal. Specific impulse has thefollowing relationship:

$I_{SP} \equiv \frac{I}{m_{p}{\mathcal{g}}} \propto \sqrt{\Delta\; h} \propto \sqrt{\frac{T_{0}}{\mathcal{M}}}$where I is the total impulse, m_(p) is the propellant mass, g isstandard gravity, Δh is the enthalpy change across the nozzle, T₀ is thestagnation temperature of the combustion products, and

is the molecular weight of the combustion products. The firstproportionality results from the assumptions that the exhaust velocityis constant and that there is isentropic flow through the nozzle, and isthe relationship that Cheetah 7.0 equilibrium code uses to calculateI_(SP) as set forth in Example 1. The second proportionality makes theassumptions that the exhaust products are an ideal and caloricallyperfect gas and shows that high chamber temperatures and low productmolecular weights improve specific impulse.

Table 2 shows that the Al—Li system has higher specific impulse thaneither neat Al or Li, regardless of the amount of Li present in Al—Li.Additionally, it is shown that hydrogen chloride formation (% Cl) issignificantly reduced with a lithium content of at least about 15% byweight.

TABLE 1 Additive Solids Theoretical AP NaNO₃ HTPB Al Li Content O/FLoading Density Additive [wt. %] [wt. %] [wt. %] [wt. %] [wt. %] [%] [−][%] [g cm⁻³] Neat Al 66.7 — 11.5 21.8 — 65.6 2.0 88.5 2.00 Neat Li 85.9—  8.2 — 5.9 41.6 6.1 91.8 1.36  Al—Li (5 wt. % Li) 65.6 — 11.1 22.2 1.267.9 1.9 88.9 1.99  Al—Li (10 wt. % Li) 64.2 — 10.8 22.5 2.5 69.9 1.889.2 1.97  Al—Li (15 wt. % Li) 63.0 — 11.0 22.0 3.9 70.2 1.7 89.0 1.95 Al—Li (20 wt. % Li) 63.0 — 11.0 20.7 5.2 70.2 1.7 89.0 1.92 *Al—Li (20wt. % Li) 61.5 — 11.7 21.4 5.4 69.6 1.6 88.3 1.91  Al—Li (25 wt. % Li)63.0 — 11.0 19.5 6.5 70.2 1.7 89.0 1.89  Al—Li (30 wt. % Li) 63.0 — 11.018.2 7.8 70.4 1.7 89.0 1.86 Neat Al, NaNO₃ 36.0 28.3 11.9 23.7 — 66.61.8 88.1 2.09

TABLE 2 Max Impulse I_(SP) Density Δh_(Chamber-Exit) T_(Chamber) Mol.Wt. Cl → HCl Cl → MCl M → MCl HPHS Additive [sec] [g-sec cm⁻³] [kJ g⁻¹][K] [kg kmol⁻¹] [%] [%] [%] [%] Neat Al 264.8 528.7 3.37 3614 27.9 98.3— —  1.7 Neat Li 263.4 358.1 3.34 3204 27.3  1.8 95.0 82.1 98.2  Al—Li(5 wt. % Li) 267.3 531.1 3.44 3658 27.8 68.5 30.0 99.3 31.5  Al—Li (10wt. % Li) 269.9 532.6 3.50 3679 27.6 34.0 65.0 98.6 66.0  Al—Li (15 wt.% Li) 271.8 529.3 3.55 3621 26.7  3.6 96.1 91.9 96.4  Al—Li (20 wt. %Li) 271.9 521.9 3.56 3553 26.2  1.9 97.8 70.3 98.1 *Al—Li (20 wt. % Li)270.7 518.3 3.52 3450 25.5  0.3 99.0 67.1 99.1  Al—Li (25 wt. % Li)272.7 515.7 3.58 3537 25.7  1.7 98.0 56.2 98.3  Al—Li (30 wt. % Li)273.4 509.7 3.60 3547 25.3  1.5 98.2 46.9 98.5 Neat Al, NaNO₃ 246.1515.0 2.91 3508 30.7 15.2 84.2 77.5 84.8

Data from Table 2 are represented in FIG. 2 and FIG. 3. The theoreticalperformance of the Al—Li system is shown in FIG. 2 as a function oflithium content. The data show that there is enough lithium present tostoichiometrically combine with the chlorine in the AP, thus yieldinggood performance as well as low hydrogen chloride content at valuesgreater than about 15% by weight lithium. It is also shown that thetheoretical performance of the Al—Li system, such as an alloy, continuesto increase with increasing lithium content. However, it is shown inFIG. 3, that above about 15%, there is an excess of lithium in the fueladditive, thus forming lithium species other than lithium chloride.While these other lithium species are shown to increase performanceabove 15%, only LiCl is desirable for an environmentally cleanerformulation.

The performance advantages of an aluminum-lithium alloy over neataluminum or neat lithium can be seen by reference to FIG. 1. The plotsin FIG. 1 were calculated in accordance with Example 1. Contour linesrepresent I_(SP) whereas transverse contour lines represent solidsloading of the propellant. FIG. 1A and FIG. 1B show impulse performancefor neat aluminum and neat lithium respectively as calculated inaccordance with Example 1. Maximum I_(SP) for each of these materials isless than that of a 20/80 lithium-aluminum alloy as seen in FIG. 1C.Performance decreases when NaNO₃ at a 56/44 weight percent AP/NaNO₃ (astandard formulation) is added to the propellant. These data aretabulated in Table 2 and it is seen that the I_(SP) performance of the20/80 alloy is 271.9 seconds whereas the neat Al and Li are only 264.8seconds and 263.4 seconds respectively whereas the NaNO₃-spiked materialonly registers 246.1 seconds. The greater the I_(SP), the more thrustthat is produced from the same amount of propellant and, therefore,increased I_(SP) is a desirable feature.

The Al—Li formulation leads to a higher specific impulse and lowercombustion product molecular weight than either of the neat parentmaterials due to a number of factors. The replacement of some aluminumwith lithium in the formulation results in the formation of some aluminaproduct (molecular weight of

$\left. {101.9\frac{\mathcal{g}}{mol}} \right)$being replaced with lithium chloride (“LiCl”) (molecular weight of

$\left. {42.4\frac{\mathcal{g}}{mol}} \right).$Furthermore, alumina will predominantly stay in the condensed phasethroughout the combustion event (boiling point of 2977° C., 1 atm),whereas LiCl will primarily remain gaseous (boiling point of 1382° C., 1atm). In some cases, the LiCl may condense out during nozzle expansion,but a chamber pressure of 68.0 atm, optimally expanded to 1.0 atm,results in a theoretical nozzle exit temperature of 2070° C., well abovethe boiling temperature of LiCl. Additionally, replacing hydrogenchloride (“HCl”) (molecular weight of

$\left. {36.5\frac{\mathcal{g}}{mol}} \right)$with LiCl only moderately increases the chlorine-based species molecularweight, but significantly lowers the global molecular weight by freeingup excess hydrogen gas (molecular weight of

$\left. {2.0\frac{\mathcal{g}}{mol}} \right).$

FIG. 3 illustrates thermochemical equilibrium calculations from Example1 for the conversion of Cl to HCl and LiCl in the process of combustingthe propellant. HCl formation shows a steady decrease towards baseline,especially as the weight percent of lithium increases above about 15%.At higher weight percentages, the amount of LiCl formed from HCl remainshigh whereas the amount of lithium products that are not LiCl increase.

According to the calculations of Example 1 and Example 2, an optimumHPHS propellant would be about 63% of AP, about 22% of an Al—Liformulation (such as an alloy) with a loading of about 83% Al and about17% Li, and about 15% of HTPB of a binder, for a solids loading of 85%.Such a solid-rocket propellant would have a specific impulse of about270.5 seconds, a hydrogen chloride content of about 2.1%, and aconversion of Li to LiCl efficiency of about 97.0% based on chlorinefrom the AP. The difference in values between such a propellant andinterpolation based on Table 2 is due to a lower solids loading in thisformulation.

In many embodiments of the invention, the oxophilic metal is aluminumand the halophilic metal is lithium and the formulation is an alloy, asthat term is usually understood by those of ordinary skill in the art.In such embodiments, the weight ratio of lithium to aluminum in thealloy is between about 14% and about 34% by weight. Further embodimentsinclude weight ratios of lithium to aluminum of between about 14% and30%, between about 14% and 24%, between about 14% and 20%, and betweenabout 16% and 18%, as well as values in between the weight ranges given.For example, separate embodiments of about 14%, 15%, 16%, 17%, 18%, 19%,or 20% are each further provided herein. When reporting values in weightpercent, the understood variability by use of the word “about” is on theorder of 1%. Thus, a weight percent of about 15% means 14% to 16%. Theuse of the word “about” is meant to modify all weight percent values setforth herein whether explicitly present or not.

The weight ratio of lithium is important to the performance of thepropellant. When the weight percent of lithium is less than 14%, thenthe amount of hydrogen chloride that is formed increases rapidly as seenin FIG. 3. Weight ratios of greater than 34% result in poor impulsedensity (total thrust per unit volume of propellant). Particularlypreferred ratios of the embodiments set forth herein are those where thephase of lithium-aluminum microcrystals in a lithium-aluminum alloy isin the simple cubic crystalline phase. Such a phase exists between about12% and about 20% by weight lithium and is particularly advantageous.The crystalline phase provides optimum performance capabilities withrespect to other phases within the acceptable weight range while alsosubstantially reducing hydrogen chloride gas formation. Such a range isalso important because as the lithium content increases over about 20%in, for example, an alloy, the amount of Li products forming, other thanthe preferred LiCl, increases substantially and free lithium is highlyreactive. Such other products may be harmful to the environment whereasLiCl is relatively benign. Thus, while lithium amounts of greater than20% may be used in a formulation with aluminum, it is preferred to use aformulation where the lithium content is in the range of between about14% and about 20% by weight, the weight range between 12% and 14%leading to a higher hydrogen chloride formation. Another embodiment iswhen substantially all of the alloy is crystalline, which occurs at aweight of about 20% lithium and 80% aluminum.

In these and other embodiments of the invention, the weight percentageof the Al—Li formulation, such as an alloy, is between about 5% andabout 40% by weight. Other embodiments include ranges between about 20%and about 40% by weight as well as between about 20% and about 30% byweight, as well as all values in between about 5% and about 40% such asabout 6%, 7%, 8%, 9%, 10%, 11%, 12%, 13%, 14%, 15%, 16%, 17%, 18%, 19%,20%, 21%, 22%, 23%, 24%, 25%, 26% 27%, 28% 29% 30%, 31%, 32%, 33%, 34%,35%, 36%, 37% 38%, or 39%.

The propellants of the invention further include an oxidizer. The amountof oxidizer in the propellant is between about 55% and about 79% byweight. Other ranges include between about 55% and about 65% by weight,between about 58% and about 65% by weight, and between about 60% andabout 64% by weight and all values in between including about 59%, 60%,61%, 67%, 63%, 64%, 65%, 66%, 67%, 68%, 69%, 70%, 71%, 72%, 73%, 74%,75%, 76%, 77%, or 78%. Oxidizers typically contain chlorine with acommon oxidizer being ammonium perchlorate.

The invention further includes propellants containing a binder. Suchbinders are often organic. Examples of binders suitable for use hereininclude hydroxyl-terminated polybutadiene (“HTPB”), carboxyl terminatedpolybutadiene (“CTBP”), Polybutadiene acrylonitrile (“PBAN”),dicyclopentadiene (“DCPD”), Polyurethane (“PU”), Plasticizednitrocellulose (“PNC”), Glycidyl Azide polymers (“GAP”), oxetanepolymers (“PolyNIMMO”), oxirane polymers (“polyGLYN”),bis-azidomethyloxetane/azideomethylmethyloxetane (“BAMO/AMMO”) orcombinations thereof. Such binders may be used to augment the fuel forcombustion. In many embodiment of the invention, the binder is presentbetween about 5% and about 25% by weight. Other ranges include betweenabout 10% and about 20% by weight. Other ranges include between about10% and about 16% and between about 11% and about 15% by weight and allvalues between about 5% and about 25% including about 6%, 7%, 8%, 9%,10%, 11%, 12%, 13%, 14%, 15%, 16%, 17%, 18%, 19%, 20%, 21%, 22%, 23%, or24%.

In certain embodiments of the invention, solid-rocket propellants areprovided wherein the Al—Li formulation is an alloy, the oxidizer isammonium perchlorate, and binder is one or more of HTPB, CTBP, PBAN,DCPD, PU, PNC, GAP, PolyNIMMO, polyGLYN, BAMO/AMMO or combinationsthereof. In such embodiments, the amount of alloy present is betweenabout 5% and about 40% by weight. Other embodiments include rangesbetween about 20% and about 40% by weight as well as between about 20%and about 30% by weight, as well as all values in between 5% and 40%such as about 6%, 7%, 8%, 9%, 10%, 11%, 12%, 13%, 14%, 15%, 16%, 17%,18%, 19%, 20%, 21%, 22%, 3%, 24%, 25%, 26%, 27%, 28%, 29%, 30%, 31%,32%, 33%, 34%, 35%, 36%, 37% 38%, or 39%. The weight ratio of lithium toaluminum in such alloys is between about 14% and about 34% by weight,including between about 14% and 30%, between about 14% and 24%, betweenabout 14% and 20%, and between about 16% and 18%, as well as about 15%,16%, 17%, 18%, 19%, or 20%. In such embodiments, the amount of ammoniumperchlorate is between about 55% and about 79% by weight. Other rangesinclude between about 55% and about 65% by weight, between about 58% andabout 65% by weight, and between about 60% and about 64% by weight andall values in between including about 56%, 57%, 59%, 60%, 61%, 67%, 63%,64%, 65%, 66%, 67%, 68%, 69%, 70%, 71%, 72%, 73%, 74%, 75%, 76%, 77%, or78%. The amount of hydroxyl-terminated polybutadiene in such embodimentsis between about 5% and about 25% by weight. Other ranges includebetween about 10% and about 20% by weight. Still other ranges includebetween about 10% and about 16% and between about 11% and about 15% byweight and all values between about 5% and about 25% including about 6%,9%, 10%, 11%, 12%, 13%, 14%, 15%, 16%, 17%, 18%, 19%, 20%, 21%, 22%,23%, or 24%.

Particle size can also influence performance characteristics in thesolid-rocket propellant of the invention. Typical particle sizes foroxidizers such as ammonium perchlorate are between about 1 micron andabout 400 microns. By “particle size” what is meant is thevolume-weighted mean particle size. In embodiments, two differentparticle sizes are used when ammonium perchlorate is the oxidizer. Insome such embodiments, the “coarse” particle size is between about 100microns 400 microns, including between about 150 microns and 250 micronsincluding all values in between such as about 200 microns. In thecontext of “particle size” the word “about” means plus or minus 10%.Thus, a value of about 100 microns means between 90 microns and 110microns. The use of the word “about” is meant to modify all particlesize measurement values set forth herein whether explicitly present ornot. In some embodiments, the “fine” particle size is between about 1micron and 75 microns, including between about 10 microns and 50microns, and further including between about 10 microns and 30 micronsincluding all values in between such as about 20 microns.

When the Al—Li formulation is an alloy, typical particle sizes arebetween about 10 microns and 200 microns including between about 10microns and about 100 microns, including between about 20 microns and 50microns, including between about 25 microns and 45 microns and values inbetween such as about 26, 27, 28, 29, 30, 31, 32, 33, 34, 35, 36, 37,38, or 39 microns. A particle-size distribution curve for Al—Li alloyprepared according to Example 3 is set forth in FIG. 4. As particle sizedecreases, the formulations become hard or brittle which isdisadvantageous from a handling and performance perspective.

The invention is also directed to chloride containing solid-fuel rocketpropellants comprising an oxidizer, an oxophilic metal-halophilic metalalloy, and a binder wherein substantially no hydrogen chloride isproduced upon combusting the fuel and wherein the alloy is capable ofscavenging hydrogen chloride. In many of the embodiments, the alloy isan aluminum-lithium alloy, the oxidizer is ammonium perchlorate and thehinder is HTPB. By “substantially no hydrogen chloride” what is meant isthat less than 5% of the hydrogen chloride that would have been producedhad a fuel been used that was not capable of scavenging hydrogenchloride to any appreciable extent such as neat Al. In otherembodiments, substantially no hydrogen chloride includes less than 3%,or 2%, or 1%. Preferable combinations of lithium to aluminum in thealloy include the ratios set forth herein such as between about 14% andabout 20% as a preferred range.

The invention is further directed to processes for making thesolid-rocket propellants of the invention described herein. In someembodiments, processes are described wherein solid-form propellantscomprising formulating a halophilic metal with an oxophilic metal toform a plurality of formulated oxophilic-halophilic metal particles andcombining the formulated oxophilic-halophilic particles with achlorine-containing oxidizer and a binder to form solid-formpropellants. Often, the formulation is an alloy and the halophilic metalis selected from lithium, sodium, potassium, rubidium, or a combinationthereof, and the oxophilic metal is selected from aluminum, beryllium,boron, magnesium, silicon, titanium, zirconium, or a combinationthereof. A typical formulation is Al—Li. In such embodiments, theoxidizer is often ammonium perchlorate and the binder ishydroxyl-terminated polybutadiene. The invention is further directed tosolid-rocket propellants so made.

Other than formulating by alloy, the invention includes other methods offormulation metals such as by mechanical activation, such as to use ballmilling to form nanocomposite, micron-scale particles, or byencapsulation wherein halophilic metal is coated with the oxophilicmetal or another agent.

The solid-rocket propellants of the invention may also include anadditive to affect performance. Such an additive may be a highexplosive, or a burn-rate modifier or a combination thereof. Examples ofhigh explosives include HMX, RDS, CL-20, TNT, or co-crystals thereof.High explosives work by increasing I_(SP). Examples of burn ratemodifiers include FeO, CuCr₂O₄, TiO₂, graphene oxide, n-butyl ferrocene,copper oxide, cobalt oxide or combinations thereof. Burn-rate modifiersaffect the rate of combustion.

The invention is further directed to processes for reducing hydrogenchloride formation during the combustion of solid-fuel rocketpropellants by preparing a solid-fuel rocket propellants of theinvention and combusting them. When compared to traditional fuels, suchas aluminum, propellants of the invention, which contains a fuel such asa lithium-aluminum formulation, can substantially reduce hydrogenchloride formation as set forth in Table 2, for example. Additionaladvantages over single-metal fuels such as aluminum is that theformulations of the invention do not form LMD during the combustionprocess. LIVID often form when aluminum is combusted which results inlong combustion residence times which is disadvantageous and may alsolead to incomplete combustion. The LMD cause both viscous and thermaldisequilibrium in a solid rocket motor (not all heat and momentum istransferred to the working gasses), resulting in motor efficiency lossesas high as 10% (i.e., two-phase flow losses).

It has been shown in the literature, that by introducing nanoscalepolymer inclusions into aluminum particles, microexplosions may beinduced at the propellant surface, although performance may becompromised. This use of microexplosions, therefore, may be able toreduce LIVID formation.

With hydrocarbon fuels, microexplosions have been widely investigated,and implemented. Because the various components within the liquid fuelsvolatilize at different temperatures, the liquid with the lowestvolatilization point will begin to outgas while other components arestill in a condensed phase. By this premise, microexplosions are causedby internal bubble nucleation and growth from within a fuel droplet(i.e., intraparticle boiling), which causes droplet fragmentation.Microexplosive liquid fuels are typically advantageous as they promoterapid fuel atomization, which can reduce residence times and increasethe completeness of combustion.

The formation of LMD from both aluminized APCP and 80/20 Al—Li APCP (seeExample 3) was investigated by using a microscopic lens withbacklighting. For the neat Al propellant, it was observed that thealuminum particles would sinter and either agglomerate or coalesce atthe surface, and then finally eject as LIVID (˜50-100 μm diameter). Incontrast, the Al—Li propellant was observed to form a bubblingmelt-layer at the surface. Most LMD that were ejected from themelt-layer would either dispersively boil or undergo rapid expansionfollowed by microexplosion into an atomized mist within a fewmillimeters of the propellant surface.

The large disparity in boiling points between aluminum (2,519° C.) andlithium (1,342° C.) allows intraparticle boiling to occur within theAl—Li alloy droplets, effectively shattering and atomizing the droplet.This microexplosive phenomenon is analogous to missive multicomponentliquid fuels that microexplode and may facilitate increased metalcombustion efficiency within a rocket motor and also decrease two-phaseflow losses. Similar particle expansion and subsequent shattering wasalso observed for neat Al—Li alloy particles that were heated via a CO₂laser at 213 W cm⁻² in air, which indicates that the chlorine is notrequired to induce shattering microexplosions with the Al—Li system,consistent with a previous study. However, the presence of chlorineappears to enhance microexplosions significantly (evident fromequivalent CO₂ laser heating of an Al—Li/AP physical mixture), possiblydue to the increase in surface heating that its presence provides (Li/CIreactions)

Example 1—Modeling

Thermoequilibrium calculations were completed using Cheetah 7.0(Lawrence Livermore National Laboratory, Edition LLNL-SM-599073)equilibrium code (JCZS product library and JCZ3 gas equation of state),assuming a chamber pressure of 6.89 MPa and ideal expansion to onestandard atmosphere. Hydroxyl-terminated polybutadiene (“HTPB”) was usedas the binder for all simulations. The following Al—Li alloycompositions were used in these simulations (wt. % Li): 0, 5, 10, 15,20, 25, 30, 100. Additionally, the Al/(56/44 wt. % AP/NaNO₃)/HTPB systemwas also investigated as a baseline scavenged propellant formulation. Atotal of 10,000 simulations were performed for each ternary system,systematically varying the oxidizer to fuel ratio (O/F) and the weightpercent of fuel additive in the total fuel (total fuel=binder+fueladditive). Post processing of these calculations was performed inMATLAB.

All Al—Li alloy fuel additives were assumed to be a physical mixture ofaluminum and lithium for all applicable equilibrium calculations, andthe binary Al—Li alloy ratio was changed with each batch of simulations.Heats of formation, Δ_(f)H, of the different Al—Li ratios actually varydue to alloy mixing and/or intermetallic formation enthalpy and areuncertain. It is recognized that this assumption yields I_(SP) valuesthat are artificially high due to increased theoretical combustiontemperatures. However, this error is expected to be minor. The LiAlintermetallic phase has the largest Δ_(f) H magnitude in the Al—Li alloyrange considered (measured to be between −9.77 to −21.8 kJ mol⁻¹), whichyields a maximum I_(SP) error of only 0.7%.

Example 2—Modeling with Scavenging

The following Equation (1) was used and evaluated as a “high performancehalide scavenger” (HPHS) metric to account for chloride scavenging:

$\begin{matrix}{{H\; P\; H\; S} = {\left( \frac{I_{SP}}{I_{{SP},\max}} \right) \times \left\lbrack {{100\%} - \left( {\%\mspace{11mu}{Cl}}\rightarrow{HCl} \right)} \right\rbrack}} & (1)\end{matrix}$where I_(SP) is the specific impulse at the current mixture ratio,I_(SP,max) is the maximum specific impulse obtainable with the ternarysystem, and % Cl→HCl is the percentage of available chlorine forming HClat the current mixture ratio and at expansion to 1 atm pressure. Thischosen performance metric places an equal weight on both specificimpulse and HCl reduction for the purpose of evaluating a cleaner, highperformance propellant formulation than otherwise available. An HPHSvalue of 100% indicates that complete MI reduction is occurring at thesystem's peak specific impulse.

Example 3—Preparation of Propellant

Two solid composite propellants were prepared using the following fueladditives: A.) neat aluminum (Alfa Aesar, −325 mesh, 99.5% purity); andB.) 80/20 wt. % Al—Li alloy (stable LiAl intermetallic phase) (SigmaAldrich). The as-received 80/20 Al—Li alloy was sieved to −325 mesh (<44μm) to be comparable with the as-received neat aluminum powder. Theparticle size distributions for both powders were determined by laserdiffraction (Malvern Mastersizer Hydro 2000 μP) using isopropyl alcoholas the dispersant medium. Surface imaging of both powders was performedby scanning electron microscopy (SEM, FEI Quanta 3D-FEG).

Imaging and particle sizing of the sieved neat aluminum (for comparison)and 80/20 Al—Li alloy powders showed that neat aluminum was nominallyequiaxed in morphology and that 80/20 Al—Li alloy had an irregularlyfaceted morphology, typically with sharp/brittle surface features. Theneat aluminum and Al—Li alloy powders had mean particle sizes(arithmetic) of 17.1 μm and 29.8 μm and volume weighted mean particlesizes (D) of 19.3 μm and 33.3 μm respectively.

The as-received 80/20 Al—Li alloy was sieved to −325 mesh (<44 μm) to becomparable with the as-received neat aluminum powder. The particle sizedistributions for both powders were determined by laser diffraction(Malvern Mastersizer Hydro 2000 μP) using isopropyl alcohol as thedispersant medium. Surface imaging of both powders was performed byscanning electron microscopy (SEM, FEI Quanta 3D-FEG).

The constituents used for the propellant formulations included: ammoniumperchlorate (ATK, 20 μm and 200 μm) and HTPB (Firefox, R45) cured withan aromatic polyisocyanate (Desmodur, E744) as the binding agent. Thefollowing formulation was used to prepare approximately 20 grams ofpropellant for each mixture:

Metal Additive: 26.8%

Coarse AP, 200 μm: 49.2%

Fine AP, 20 μm: 12.3%

HTPB (11.5% curative): 11.7%

For comparison with theoretical performance predictions in FIG. 1A andFIG. 1C, these ratios correspond to an O/F of 1.60, a fuel additive wt.% of 69.6%, and a solids loading of 88.3%. No incompatibilities wereobserved with the aromatic polyisocyanate curative, though the workingtime of the wetted propellant was short (approximately 30 minutes).

Propellant constituents were resonant mixed (Resodyn LabRAM resonantmixer) in a 60 mL container (McMaster-Carr 42905T23) for 10 min at 90%intensity. Strands were then packed into 5.8 mm diameter cylindricalmolds and cured in air for approximately 3 days at room temperature. Theburning characteristics of the propellants were investigated using acolor high-speed camera (Vision Research, Phantom v7.3) at 1000 fps in avented fume hood.

Example 4 HCl Detection—Experimental Setup

Both neat aluminum and Al—Li alloy propellants were burned in a closed130 mL stainless steel Parr cell filled with 50 ml distilled water(i.e., wet bomb experiments), since this method has been used inprevious studies. A schematic of the wet bomb experiment is shown inFIG. 8. Propellant strands (50.0±0.2 mg) were burned in argon at 2.5MPa. The pH of the distilled water was measured with a digital pH meter(Omega, PHH-37) before and after each wet bomb experiment. Theexperiment was repeated three times for each propellant type.

Neat powders of AP (200 μm, Firefox) and stirred mixtures of 30/70 wt. %metal/AP were also analyzed using simultaneous differential scanningcalorimetry/thermogravimetric analysis (DSC/TGA) (Netzsch Jupiter STA449F1) with online gas analysis via mass spectrometry (MS, Netzsch Aeolos)and Fourier transform infrared absorption (FTIR, Bruker Tensor 37). The30/70 wt. % ratio of metal/AP was chosen, as it is similar to themetal/AP ratio of the experimental propellants. Experiments wereconducted on 3.1 mg samples of AP and 4.1 mg samples of metal/AP suchthat the total amount of AP contained within each sample was similar.

Prior to starting experiments, the DSC/TGA instrument's platinum furnacewas evacuated and backfilled with purge gas three times. The sampleswere heated from room temperature to 700° C. at a heating rate of 20° C.min⁻¹ under a steady flow of 40 mL min⁻¹ of ultra-high purity argon(99.999 vol. %). The instrument exhaust was coupled to FTIR and MSinstrumentation using a heated (200° C.) manifold and silicacapillaries. The capillary was interfaced with the FTIR instrument'sliquid nitrogen-cooled MCT detector via a heated (200° C.) Bruker TGA-IRlight pipe. During DSC/TGA experiments, MS and IR data was recorded attemperature increments of approximately 3° C. with a mass range of 10 to100 and IR spectral resolution of 2 cm⁻¹, respectively. Data waspost-processed using the Netzsch Proteus software, NISTMS, Bruker Opussoftware packages, and NIST/EPA MS and FTIR libraries.

Example 5—Performance and HCl Detection—Results

The reduction in hydrogen chloride for Al—Li based propellants wasevaluated using wet bomb combustion experiments and DSC/TGA coupled withsimultaneous mass spectrometry and FTIR. For the wet bomb combustionexperiments, a pH of 2.10±0.04 was measured for the aluminum propellantand a pH of 2.71±0.08 for the Al—Li propellant. This results in a75.5±4.8% reduction of [H⁺], which is proportional to the relativechange in HCl concentration. The theoretical HCl reduction between thetwo propellants is 91% at this mixture ratio and pressure. It ispossible that the small scale of the experiment inhibited completecombustion of the propellants due to quenching of the plume against thecold chamber walls and water, causing the theoretical and measuredvalues to vary. However, the HCl reduction is still substantial. Thisexperiment has been shown to be representative of other methods employedto quantify HCl content within rocket motor plumes and similar losses tocomplete combustion have been observed.

The DSC/TGA-MS traces for neat AP in FIG. 5 suggests that there is anendothermic crystallographic phase change from orthorhombic to cubic atroughly 240° C. (no gas evolution detected), which is followed by abroad exothermic decomposition at roughly 300° C. (evolution of N₂O, NO,and H₂O detected). This broad decomposition is followed by a rapidexothermic decomposition, with an onset at approximately 410° C.(evolution of HCl, N₂O, NO, H₂O, and O₂ detected), resulting in a totalmass loss of 75%. A similar trend is apparent with the addition ofaluminum as shown in FIG. 6, indicating that the presence of aluminumhas little effect on the heating rate decomposition mechanisms of AP.

When AP is in the presence of 80/20 Al—Li as illustrated in FIG. 7,exothermic decomposition of AP is reduced to approximately 180° C.(evolution of NH₃, NO, N₂O, and H₂O detected), coincident with themelting point of lithium. The exothermic decomposition is immediatelyfollowed by a very weak endothermic phase change at roughly 240° C. (nonew gas evolution detected). A second, rapid exothermic decompositiononsets at 360° C. (evolution of NH₃, NO, N₂O, H₂O, and O₂ detected),followed by a third exothermic decomposition onset at approximately 420°C. (evolution of NH₃ and O₂ detected). The exothermic processes werethen followed by three distinct endothermic events at roughly 540° C.,600° C., and 660° C. (no new gas evolution detected). The last endothermat 660° C. corresponds to the melting temperature of neat aluminum; themelting point of 80/20 Al—Li alloy is 695° C., suggesting that lithiumwas extracted from the Al—Li phase prior to 660° C., likely at the 180°C. exotherm. The roughly 600° C. endotherm may correspond to either theeutectic melting point in the Al—Li phase diagram (596° C., at 26 at %lithium) or the melting point of LiCl (605° C.), which would indicatethat LiCl formation had occurred prior to that temperature at one ormore of the previous exothermic events. No HCl evolution was detected inthe Al—Li/AP system. Furthermore, the presence of Al—Li alloy was ableto reduce AP decomposition significantly from roughly 300° C., to 180°C. (well below the AP crystallographic phase change), likely due to LiClformation. Moreover, mass spectral data show no evidence of HCl whereasHCl was detected in both HG. 5 and FIG. 6.

FTIR measurements in FIG. 9 confirmed that neat AP yields a strong HClvibration band between 3100-2600 cm⁻¹ at approximately 450° C. However,from FTIR measurements, it was also apparent that there is low-levelformation of HCl at roughly 320° C., coincident with the first exothermshown in FIG. 5. Virtually identical HCl vibrational bands at the sametemperatures were evident with Al/AP FTIR measurements, again indicatingthat the presence of aluminum has little, if any, effect on the heatingrate HCl evolution from AP decomposition. The FTIR measurements between3100-2600 cm⁻¹ from Al—Li/AP indicates that there are some broadabsorption features above 2800 cm⁻¹ near 180° C. and 360° C. (coincidentwith Al—Li/AP exothermic peaks), but no strong HCl vibrational bands aredetected at any temperatures as were detected with neat AP. Closeobservation of the Al—Li/AP plots indicates that there may be somelow-level HCl vibrational bands evident above 360° C., but absorptionsignal levels are near the instrument noise floor, thus HCl presencecannot be confirmed. However, the data are consistent with only a smallamount of HCl.

The combustion characteristics varied greatly between the Al and Al—Lipropellants. The propellant testing was performed in air at roomtemperature and atmospheric pressure. Neat aluminized propellant burningsurfaces were dominated by large, bright aluminum droplets lifting offof the propellant surface. Droplets were suspended briefly above theburning surface, and then gently fell. The Al—Li propellant wascharacterized by a large, bright magenta flame and a relatively darkpropellant surface. The magenta flame is indicative of LiCl emission.Very few coarse droplets were observed ejecting from the Al—Li surface.Such droplets because they are undesirable in rocket propellants cancause two-phase flow losses.

Example 6

In another embodiment, the Al—Li alloy may be cryomilled in order tosafely reduce the particle size of the as-received powder until it is ina favorable range. The cryomilled Al—Li alloy can then be sieved to adesired particle size distribution in preparation for propellant mixing.

The materials used for composite propellant formulations can include:neat aluminum powder, Al—Li alloy, ammonium perchlorate, and apolybutadiene based binding agent. Propellant can be mixed according tothe various mixture ratios outlined in Table 1. Increased feasibility inrocket motor systems (e.g., castability) may be realized by increasingbinder contents in Table 1 to be between about 5% to about 25%, with themetal-to-AP ratios kept relatively constant. Mixing of the propellantcan be accomplished via resonant mixing, shaker mixing, or physicalmixing/stirring.

Example 7—Propellant

A rocket propellant of the following combination was prepared:

AP: 62.963%

Al—Li Alloy (80/20 wt. %): 18.594%

Aluminum: 3.443%

HTPB binder: 15.000%

Where the materials were sourced as set forth in Example 3. These werehand mixed with a spatula to prepare the propellant.

Example 8—Microexplosion Investigations

Solid form rocket propellants made according to Example 3 were used inthis investigation. The burning characteristics of the propellantsurfaces were investigated using a color high speed camera (VisionResearch, Phantom v7.3) at 9900 fps and using a long distancemicroscopic optic (Infinity Photo-Optical, K2 Lens, CF2 objective).Imaging was performed in air and the propellant strands were backlitwith a green light emitting diode (LED, model CREEXPE2-COL-X with amodel 10003 20 mm narrow spot LED lens). The LEI) light source wasexpanded and aligned behind the propellants strands using two concavespherical aluminum mirrors (Edmond Optics, 152 cm focal length).

Laser ignition of loose powder samples was performed using a CO₂ laser(Coherent GEM 100 A, 10.6 μm wavelength). The beam was focused using aZnSe plano-convex focusing lens (ThorLabs, model LA7270-F), yielding aspot size diameter of 1.29 mm. Approximately 2 mg of powder (neat Al—Lior Al—Li/AP) was placed on a ceramic tile (OZM Research, Part NumberBFST-Pt-100) and arranged into a thin row (<0.5 mm). Particle combustionwas observed using high speed videography (Vision Research, Phantomv7.3) at 4000 fps and using a 200 mm macro lens (Nikon f/4 AF-DMicro-NIKKOR).

The invention claimed is:
 1. A solid-rocket propellant comprising analuminum-lithium alloy, aluminum, an oxidizer, and a binder wherein theratio of lithium to aluminum in the alloy is between 17% and about 34%by weight.
 2. The solid-rocket propellant of claim 1, wherein the alloyis crystalline.
 3. The solid-rocket propellant of claim 2, wherein thecrystalline phase of the alloy is simple cubic.
 4. The solid-rocketpropellant of claim 1, wherein the amount of aluminum-lithium alloy inthe propellant is between about 5% and about 40% by weight, the amountof oxidizer is between about 55% and about 79% by weight, and the amountof binder is between about 5% and about 25% by weight.
 5. Thesolid-rocket propellant of claim 4, wherein the amount ofaluminum-lithium alloy in the propellant is between about 20% and about40% by weight, the amount of oxidizer is between about 55% and about 65%by weight, and the amount of binder is from about 8% to about 15%. 6.The solid-rocket propellant of claim 1, further comprising an additive.7. The solid-rocket propellant of claim 6, wherein the additive is ahigh explosive, a catalyst, a burn-rate modifier, or a combinationthereof.
 8. The solid-rocket propellant of claim 1, wherein the oxidizeris ammonium perchlorate, and the binder is hydroxyl-terminatedpolybutadiene, carboxyl terminated polybutadiene, Polybutadieneacrylonitrile, dicyclopentadiene, Polyurethane, Plasticizednitrocellulose, Glycidyl Azide polymers, oxetane polymers, oxiranepolymers, bis-azidomethyloxetane/azideomethylmethyloxetane orcombinations thereof.
 9. The solid-rocket propellant of claim 5, whereinthe oxidizer is ammonium perchlorate, and the binder ishydroxyl-terminated polybutadiene, carboxyl terminated polybutadiene,Polybutadiene acrylonitrile, dicyclopentadiene, Polyurethane,Plasticized nitrocellulose, Glycidyl Azide polymers, oxetane polymers,oxirane polymers, bis-azidomethyloxetane/azideomethylmethyloxetane orcombinations thereof and wherein the alloy is crystalline.
 10. Thesolid-rocket propellant of claim 1, wherein the ratio of lithium toaluminum in the propellant is between about 14% and 24% by weight. 11.The solid-rocket propellant of claim 10, wherein the ratio of lithium toaluminum is between about 16% and 18% by weight and the alloy iscrystalline.
 12. The solid-rocket propellant of claim 11, wherein theoxidizer is ammonium perchlorate and the binder is hydroxyl-terminatedpolybutadiene, carboxyl terminated polybutadiene, Polybutadieneacrylonitrile, dicyclopentadiene, Polyurethane, Plasticizednitrocellulose, Glycidyl Azide polymers, oxetane polymers, oxiranepolymers, bis-azidomethyloxetane/azideomethylmethyloxetane orcombinations thereof.
 13. The solid-rocket propellant of claim 1,wherein the ratio of lithium to aluminum is between about 19% and about21% by weight and the alloy is crystalline.
 14. The solid-rocketpropellant of claim 13, wherein the oxidizer is ammonium perchlorate andthe binder is hydroxyl-terminated polybutadiene, carboxyl terminatedpolybutadiene, Polybutadiene acrylonitrile, dicyclopentadiene,Polyurethane, Plasticized nitrocellulose, Glycidyl Azide polymers,oxetane polymers, oxirane polymers,bis-azidomethyloxetane/azideomethylmethyloxetane or combinationsthereof.
 15. A solid-rocket propellant comprising an aluminum-lithiumalloy, wherein the ratio of lithium to aluminum in the alloy is between17% and about 34% by weight, an oxidizer, and a binder.
 16. Thesolid-rocket propellant of claim 15, further comprising neat aluminum.17. The solid-rocket propellant of claim 16, wherein the alloy iscrystalline.
 18. The solid-rocket propellant of claim 17, wherein thecrystalline phase of the alloy is simple cubic.
 19. The solid-rocketpropellant of claim 15, wherein the amount of aluminum-lithium alloy inthe propellant is between about 5% and about 40% by weight, the amountof oxidizer is between about 55% and about 79% by weight, and the amountof binder is between about 5% and about 25% by weight.
 20. Thesolid-rocket propellant of claim 19, wherein the amount ofaluminum-lithium alloy in the propellant is between about 20% and about40% by weight, the amount of oxidizer is between about 55% and about 65%by weight, and the amount of binder is from about 8% to about 15%. 21.The solid-rocket propellant of claim 15, further comprising an additive.22. The solid-rocket propellant of claim 21, wherein the additive is ahigh explosive, a catalyst, a burn-rate modifier, or a combinationthereof.
 23. The solid-rocket propellant of claim 15, wherein theoxidizer is ammonium perchlorate, and the binder is hydroxyl-terminatedpolybutadiene, carboxyl terminated polybutadiene, Polybutadieneacrylonitrile, dicyclopentadiene, Polyurethane, Plasticizednitrocellulose, Glycidyl Azide polymers, oxetane polymers, oxiranepolymers, bis-azidomethyloxetane/azideomethylmethyloxetane orcombinations thereof.
 24. The solid-rocket propellant of claim 20,wherein the oxidizer is ammonium perchlorate, and the binder ishydroxyl-terminated polybutadiene, carboxyl terminated polybutadiene,Polybutadiene acrylonitrile, dicyclopentadiene, Polyurethane,Plasticized nitrocellulose, Glycidyl Azide polymers, oxetane polymers,oxirane polymers, bis-azidomethyloxetane/azideomethylmethyloxetane orcombinations thereof and wherein the alloy is crystalline.
 25. Thesolid-rocket propellant of claim 16, wherein the ratio of lithium toaluminum in the propellant is between about 14% and 24% by weight. 26.The solid-rocket propellant of claim 25, wherein the ratio of lithium toaluminum is between about 16% and 18% by weight and the alloy iscrystalline.
 27. The solid-rocket propellant of claim 26, wherein theoxidizer is ammonium perchlorate and the binder is hydroxyl-terminatedpolybutadiene, carboxyl terminated polybutadiene, Polybutadieneacrylonitrile, dicyclopentadiene, Polyurethane, Plasticizednitrocellulose, Glycidyl Azide polymers, oxetane polymers, oxiranepolymers, bis-azidomethyloxetane/azideomethylmethyloxetane orcombinations thereof.
 28. The solid-rocket propellant of claim 16,wherein the ratio of lithium to aluminum is between about 19% and about21% by weight and the alloy is crystalline.
 29. The solid-rocketpropellant of claim 28, wherein the oxidizer is ammonium perchlorate andthe binder is hydroxyl-terminated polybutadiene, carboxyl terminatedpolybutadiene, Polybutadiene acrylonitrile, dicyclopentadiene,Polyurethane, Plasticized nitrocellulose, Glycidyl Azide polymers,oxetane polymers, oxirane polymers,bis-azidomethyloxetane/azideomethylmethyloxetane or combinationsthereof.